XFOIL Version 6.94 Calculated polar for: NACA4412c 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.070 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0445 0.03416 0.02465 -0.0512 0.9999 0.1705 -2.750 -0.0187 0.03372 0.02417 -0.0511 0.9999 0.1806 -2.500 0.0078 0.03323 0.02364 -0.0512 0.9999 0.1926 -2.250 0.0334 0.03317 0.02346 -0.0511 0.9999 0.2069 -2.000 0.0973 0.03111 0.02163 -0.0580 0.9884 0.2366 -1.750 0.1701 0.02864 0.01939 -0.0659 0.9704 0.2841 -1.500 0.2469 0.02544 0.01703 -0.0744 0.9492 0.4088 -1.250 0.3029 0.02127 0.01421 -0.0760 0.9079 1.0001 -1.000 0.3492 0.01975 0.01217 -0.0761 0.7897 1.0001 -0.750 0.3720 0.02039 0.01114 -0.0722 0.5300 1.0001 -0.500 0.3967 0.02143 0.01128 -0.0714 0.4563 1.0001 0.000 0.4552 0.02272 0.01172 -0.0718 0.4148 1.0001 0.250 0.4853 0.02329 0.01200 -0.0722 0.4037 1.0001 0.500 0.5153 0.02389 0.01231 -0.0725 0.3943 1.0001 0.750 0.5454 0.02452 0.01270 -0.0729 0.3865 1.0001 1.000 0.5756 0.02512 0.01313 -0.0733 0.3790 1.0001 1.250 0.6056 0.02588 0.01362 -0.0738 0.3729 1.0001 1.500 0.6360 0.02662 0.01423 -0.0743 0.3683 1.0001 1.750 0.6662 0.02731 0.01487 -0.0748 0.3640 1.0001 2.000 0.6963 0.02807 0.01556 -0.0753 0.3599 1.0001 2.250 0.7263 0.02891 0.01630 -0.0758 0.3562 1.0001 2.500 0.7562 0.02991 0.01715 -0.0764 0.3525 1.0001 2.750 0.7858 0.03091 0.01811 -0.0769 0.3491 1.0001 3.000 0.8152 0.03171 0.01901 -0.0774 0.3454 1.0001 3.250 0.8443 0.03263 0.01997 -0.0779 0.3420 1.0001 3.500 0.8735 0.03364 0.02102 -0.0784 0.3392 1.0001 3.750 0.9023 0.03472 0.02213 -0.0789 0.3367 1.0001 4.000 0.9310 0.03591 0.02333 -0.0793 0.3344 1.0001 4.250 0.9593 0.03728 0.02470 -0.0798 0.3320 1.0001 4.500 0.9868 0.03886 0.02635 -0.0803 0.3296 1.0001 4.750 1.0134 0.04002 0.02775 -0.0806 0.3271 1.0001 5.000 1.0396 0.04137 0.02932 -0.0809 0.3246 1.0001